The present invention relates to an axial flow gas turbine engine in which fuel burners are connected by transition ducts to a turbine inlet. The transition ducts are flexibly supported within the engine to tolerate the high temperatures of the propulsive gases transmitted from the burners through the transition ducts to the inlet.
In all gas turbine engines, burners in the combustion section receive and react fuel and compressed air to produce propulsive gases which are transmitted from the burners to the turbine section and propel one or more turbine wheels. The turbine wheels drive either the engine compressor or a power shaft from which the engine power is taken. The burners, sometimes referred to as burner cans, are usually cylindrical in shape and are distributed circumferentially about the engine axis to receive compressed air from the engine compressor and fuel from the fuel system. The fuel and air are mixed within the burners and are ignited in a combustion process which produces the propulsive gases. The gases are transmitted from the burners to the turbine inlet through transition ducts which become relatively hot due to the high temperature of the propulsive gases. Special provisions are sometimes made at the turbine inlet in order to cool the hot components in that region.
U.S. Pat. No. 2,743,579 discloses a gas turbine engine having burners including a transition portion which varies in shape along its axial length from a generally circular form at the forward end to an arcuate form at its aft end so that a series of such transition portions may mate with an annular turbine inlet. The arcuate ends are supported by inner and outer rings and a plurality of struts extending between the rings, each ring being secured to the engine frame. To alleviate thermal stresses that might be generated in the support rings, and inlet guide vanes, cooling air from the compressor is directed into the transition portion over the rings and through the guide vanes.
The present invention has as a general object the provision of the improved structure for supporting transition ducts at the turbine inlet with greater compliance in the presence of the high temperature propulsive gases. It is also an object of the invention to provide bulkhead sealing between the combustion and turbine sections and thereby impede the leakage of compressed air directly into the turbine inlet.